The invention relates to a method and apparatus for determining the position and orientation of a missile, particularly a space missile.
Inertial navigation systems and sometimes supplementary sensor systems such as barometric sensors, are generally used to determine the position and orientation of spacecraft, including space missiles (such as booster rockets and their payloads), satellites, and other exo- and transatmospheric flying devices. “Orientation” refers to the location/orientation of the space missile in space, t.e., the angles of roll, pitch, and yaw. Although such systems are independent of external devices, and in particular to external signals, they entail high costs. In addition, they do not provide information concerning the absolute position of a space missile.
In U.S. Pat. No. 6,535,801 B1, the absolute position and orientation of a satellite in a geosynchronous orbit are determined using global positioning system (GPS) signals. Because GPS satellites are oriented such that their principal direction of radiation for GPS signals is pointed toward the earth, in geosynchronous orbits (which are situated above the orbits of GPS satellites) on average only one GPS signal (or even none at all, but no more than three GPS signals) is received, which is generally insufficient for a reliable determination of the absolute position and orientation of a geosynchronous satellite. It was therefore proposed to process GPS signals transmitted by the GPS satellites in secondary directions of radiation as well; however, these signals are very noisy.
According to U.S. Pat. No. 6,535,801, such processing may be enabled, for example, by using a pseudolite stationed on the earth, the signal from which is continuously received by the geosynchronous satellite and used for a highly accurate GPS time standard in the receiver of the geosynchronous satellite. A disadvantage of this method, however, is that the GPS satellites signals which are used are under the control of the United States Department of Defense, and therefore may be switched off or become degraded at any time.
One object of the present invention is to provide a system and a method for determining the position of a missile, in particular a space missile, which enables the absolute position to be determined and which is as independent as possible from GPS signals.
This and other objects and advantages are achieved by the method and apparatus according to the invention, in which at least one pseudolite (short for “pseudosatellite”) station on the earth's surface transmits a navigation signal into space (i.e., the near-earth space outside the earth's atmosphere and also interplanetary space and space navigable by space missiles) which uniquely identifies the station. The navigation signal transmitted by the psetidolite station may then be received by a missile (for example a rocket) and used to determine the absolute position of the missile, either i) solely on the basis of received navigation signals from multiple pseudolite stations, ii) in interaction with inertial navigation systems that are present, or iii) with other navigation signals from satellites. By use of such a pseudolite station the above-described disadvantages of position determination by means of satellites may be largely avoided, in particular problems in reception of navigation signals at altitudes greater than the orbits of navigation satellites.
Heretofore, pseudolite stations have been used to improve navigational accuracy on the earth's surface, and have therefore transmitted their navigation signals only in the region of the earth's surface, to enable optimal reception by navigation signal receivers on the earth's surface or in the vicinity thereof. Transmitting a navigation signal from a pseudolite station into space according to the present invention provides the possibility for determining the absolute position of missiles also at very high altitudes, in particular for space missiles such as rockets, or also geosynchronous satellites.
The invention may be used alone or in addition to inertial navigation systems present in missiles to enable a particularly accurate and, above all, an absolute determination of the missile's position and orientation. For example, operators of rocket launching pads may utilize the invention to be independent from operators of satellite navigation systems and to maintain complete control over the system for determining the absolute position and orientation of rockets. The invention is suitable for any type of missile, in particular for applications in which conventional navigation systems such as GPS or the Global Navigation Satellite System (GLONASS or GNSS) are too imprecise, are unsuitable due to high flight altitude, or are too unreliable. Lastly, the invention may also be implemented in missiles more economically than is the case for current avionics for position and orientation determination.
According to one embodiment, the invention relates to a system for determining the position of a missile, in particular a space missile, comprising a ground segment having at least one pseudolite station which is located on the earth's surface and which transmits a navigation signal into space which uniquely identifies the pseudolite station, and a user segment having at least one sensor, situated in the missile, which is designed to receive the navigation signal of the at least one pseudolite station and to determine its absolute position in space on the basis of the received navigation signal. By use of such a system, the position of a missile may be determined largely independently of satellite navigation systems, but at a minimum the accuracy of the position determination may be significantly improved by use of satellite navigation signals. The system may also be used as a fall-back solution for the position determination, for example if an inertial navigation system of a missile fails, or satellite-based navigation is unusable due to poor reception of navigation signals.
To increase the accuracy of the position determination and in particular to enable the most accurate position determination possible at various altitudes of the missile and flight phases, the ground segment may include multiple pseudolite stations which are configured on the earth's surface according to a specified geometry, in such a way that navigation signals from at least five different pseudolite stations may be received during all phases of flight of the missile. In this manner the position may be determined with an accuracy that is sufficient in particular for critical missions such as launching and flight of a booster rocket for satellites, and which is more reliable, at least in part, than position determination based on satellite navigation signals.
Because particularly high accuracy is required for the position determination, above all in critical flight phases such as the launch, the pseudolite stations may also be configured in such a way that navigation signals may be received from a maximum number of pseudolite stations during the launch of the missile from the earth's surface. For example, multiple pseudolite stations may be provided at short distances from one another in the region of a rocket launching pad so that there is a high probability that a launching booster rocket will receive navigation signals from a large number of pseudolite stations and is able to determine its absolute position in space with a high degree of accuracy.
According to a further embodiment, the invention relates to a pseudolite station for use in a system according to the invention, wherein the psuedolite station comprises a signal generator which generates a navigation signal that uniquely identifies the pseudolite station, and transmits the navigation signal via a code multiplexing process, and a transmitter antenna for transmitting the code multiplexing signal, which signal generator uses a code uniquely associated with the pseudolite station as a spreading code.
In principle, a pseudolite station may have a design that is similar to a navigation satellite, except that it occupies a fixed, specified position on the earth's surface and transmits its navigation signals into space, not to the earth's surface as is the case with a satellite.
In particular, a control and monitoring unit may be provided which is designed to control the generation of navigation signals and to monitor the transmission of navigation signals in coordination with data for a time standard.
To ensure that multiple pseudolite stations transmit their navigation signals based on the most synchronous time standard possible, a monitor station for monitoring the pseudolite station may be provided to receive navigation signals from satellites, and to generate synchronization signals for the control and monitoring unit of the pseudolite station from the received navigation signals. Due to the curvature of the earth, however, centralized synchronization of the pseudolite station is difficult if not impossible. It is therefore advantageous to synchronize satellite navigation signals so that the signals may be easily received by the pseudolite stations.
The monitor station may be designed in particular to extract time information from navigation signals received from satellites, and to use this information to generate synchronization signals. In other words, the monitor station of a pseudolite station may make use of external satellite navigation signals as a time standard.
Alternatively or additionally, the monitor station may be designed to obtain time information by means of a two-way time transfer process, and to use this information to generate synchronization signals.
A pseudolite station may also have a high-stability oscillator reference as a local time standard, by means of which to a certain extent the pseudolite station may operate autonomously; that is, in particular, independently of external signals such as satellite navigation signals.
According to a further embodiment, the invention relates to a sensor for use in a system according to the invention, which sensor comprises at least one antenna and at least one receiver for receiving a navigation signal from at least one pseudolite station, and processing and computation means for processing the received navigation signal and computing the absolute position and orientation of the sensor from the processed navigation signal. Such a sensor may be, for example, a unit which may be installed in various missiles (for example booster rockets and their payloads, satellites, and other exo- and transatmospheric flying devices), to allow the absolute position of the missile to be determined according to the invention.
The at least one antenna and the at least one receiver may also be designed to receive navigation signals from satellites. In this manner the accuracy of the position determination may be further improved, or sometimes even increased. Received satellite navigation signals may, for example, be entered directly into the position determination, or may also be used to check the plausibility of a position which has already been determined on the basis of the navigation signals from pseudolite stations.
A further improvement in the position determination may be achieved by providing the processing and computation means to process data from an inertial navigation system of a missile, and to take the data from the inertial navigation system into account in computing the absolute position and orientation of the sensor from the processed navigation signals. In this regard, an inertial navigation system is understood to mean all components which are used for determining the relative position of a missile, i.e., also position sensors and barometric sensors, for example.
The processing and computation means in particular have a digital signal processor which is programmed to reconstruct and decode navigation signal data from the received navigation signals.
In one preferred embodiment, the processing and computation means also has a central processing unit which is programmed to compute the absolute position and orientation of the sensor from the navigation signal data generated by the digital signal processor, taking into account the position data from at least one pseudolite station. In this case the algorithms, which are used for position determination and in principle correspond to the algorithms for position determination using satellite navigation signals, are essentially implemented in software which is run by the central processing unit. However, the algorithms may also be implemented in hardware in order to achieve, for example, the most rapid position determination possible.
The processing and computation means may also have a data memory in which position data from at least one pseudolite station together with the unique code for the pseudolite station are stored. The central processing unit is then programmed in such a way that it is able to read from the data memory the position data associated with the pseudolite station on the basis of navigation signal data from a pseudolite station, and to compute the absolute position and orientation of the sensor. In this case it is not necessary, in principle, for the pseudolite stations to transmit position data in their navigation signals. In principle, the unique code for a given pseudolite station is sufficient to determine in the sensor the positions of the pseudolite station on the basis of the information stored in the data memory, and to use this information for the position determination. However, if the position data are transmitted with the navigation signals, for poor or faulty signal reception, for example, in which only the pseudolite station transmitting the signal can be ascertained, position determination is still performed, since the corresponding position data from the pseudolite station are locally present in the sensor.
Lastly, according to one embodiment the invention relates to a method for determining the position of a missile, in particular a space missile, wherein at least one pseudolite station of a ground segment which is located on the earth's surface transmits a navigation signal into space which uniquely identifies the pseudolite station, and at least one sensor for a user segment, provided in the missile, receives the navigation signal from the at least one pseudolite station and determines its absolute position in space on the basis of the received navigation signal.
In determining its absolute position in space, the sensor may also use at least one navigation signal from a satellite.
In determining its absolute position in space, the sensor may also use additional data from an inertial navigation system for the missile.
Other objects, advantages and novel features of the present invention will become apparent from the following detailed description of the invention when considered in conjunction with the accompanying drawings.